The present invention relates to an axial compressor for a gas turbine, a gas turbine having the axial compressor, and an aircraft engine having the gas turbine.
Rotor blades and guide vanes of gas turbine axial compressors have a blade or vane element arranged in a flow duct of the axial compressor for redirecting the flow, said blade or vane element having a leading edge and a trailing edge, which are joined to each other through a pressure side and a suction side.
Between a tip of a blade or vane and the duct wall of the flow duct that lies radially opposite to it, a (radial) gap is normally present owing to relative rotation and, as a result of it, any flow redirected by the blade or vane is disrupted in a detrimental way.
This gap can vary. In particular, it can increase in size over the operating time of the axial compressor or—on account of the reduced centrifugal force—at lower speeds of rotation.
This increase in the gap detrimentally enhances the disruption induced by the gap and thereby deteriorates the (aerodynamic) performance of the blade or vane and hence of the axial compressor, in particular the efficiency of the blade or vane or the efficiency of the axial compressor, and/or its surge margin or its (pumping) stability. A deteriorated surge margin needs to be dealt with in the design and leads, in turn, to a (further) deterioration in the efficiency.
It is therefore desirable to reduce the deterioration in the performance and/or (pumping) stability in the case of gaps that are larger or are becoming larger and to reduce the gap sensitivity.